In an aircraft gas turbine (jet) engine, air is drawn into the front of the engine, compressed by a shaft-mounted compressor, and mixed with fuel. The mixture is combusted, and the resulting hot combustion gases are passed through a turbine mounted on the same shaft. The flow of gas turns the turbine by contacting an airfoil portion of the turbine blade, which turns the shaft and provides power to the compressor. The hot exhaust gases flow from the back of the engine, driving it and the aircraft forward. There may additionally be a bypass fan that forces air around the center core of the engine, driven by a shaft extending from the turbine section.
In service, the components of the turbine section of the engine are repeatedly thermally cycled between lower and higher temperatures. The thermal cycles may be between room temperature and the operating temperature when the engine is started, operated, and shut down, or between an intermediate temperature and a higher operating temperature when the engine power output is changed between lower and higher power settings. Some of the components are mechanically loaded during the thermal cycling.
One of the components that is subjected to the greatest temperature extremes and the greatest mechanical loadings during thermal cycling is the turbine blade. The turbine blade has a complex shape, including an attachment section, an airfoil with a thin tip portion, and a platform. The airfoil is in the direct flow of the hot combustion gases. The turbine blade is typically made of an alloy that has good high temperature mechanical properties.
As a result of its material of construction and the nature of its thermal and mechanical loadings during service, the turbine blade, particularly the airfoil and airfoil tip, are subject to thermo mechanical fatigue (TMF). The combination of thermally induced stresses, mechanically induced stresses, and the temperature extremes can combine to cause the formation and propagation of cracks in the airfoil of the turbine blade, particularly in its airfoil tip, that lead to premature failure of the turbine blade. Because the turbine blade rotates at up to 20,000 rpm, such a premature failure may lead to an unplanned removal from service or major damage of the engine.
It is therefore important to evaluate the performance of the turbine blade with regard to its material of construction and the loading conditions to ensure that the turbine blade can withstand the thermal mechanical fatigue and will not fail prematurely. There are a variety of testing procedures used, including full scale tests of the turbine blades, tests of proxy specimens such as cylindrical test bars which are used in the hope that their performance will reflect the performance of the turbine blades, and accelerated tests using special test equipment. All of the available testing approaches have shortcomings. The full-scale tests are expensive, the tests of proxy specimens do not necessarily reflect the performance of the turbine blade, and accelerated tests are not a good indicators of actual service performance and do not permit a realistic evaluation of the materials and conditions. Stresses imposed in other simulative tests fail to accurately account for the complex effects of the test material in determining the stress that results due to a number of factors such as thermal expansion coefficient, elastic modulus, emissivity, thermal conductivity, and yield strength, all of which vary with temperature.
There remains a need for an improved approach to evaluating components of gas turbine engines, as well as articles subjected to similar conditions, to evaluate their thermal mechanical fatigue performance. The present invention fulfills this need, and further provides related advantages.